This invention relates to a gas turbine engine and a method for controlling a gas flow rate in an axial flow gas turbine engine. In particular, the invention relates to a gas turbine engine arranged for propulsion of an aircraft.
The fuel consumption of gas turbine engines is normally higher at part power than at full power due to the lower pressure and temperature ratios (between compressed air and inlet air) at part power operation. Considerable savings in fuel consumption (around 5% for a commercial turbofan aircraft engine) are potentially possible to achieve if the gas turbine engine can be operated with a higher pressure and temperature at part power. It is known that application of a variable cycle to an axial flow gas turbine engine, such as an aircraft turbo-jet engine, is useful for this purpose. In short, to apply a variable cycle to a gas turbine engine means that the air mass flow and the pressure ratio of a certain engine can be adapted to different operation modes. The effects of variable cycle gas turbine engines are, for instance, described in U.S. Pat. Nos. 4,287,715, 5,806,303 and Lundbladh, A. and Avellan, R., Potential of variable cycle engines for subsonic air transport, ISABE 2007-1156, 2007.
There are various ways of achieving a variable cycle turbofan engine. The two most common ways is to provide a variable fan flow or a variable core flow. Of these, the variable fan flow is primarily useful for a supersonic aircraft.
Most variable cycle gas turbine engines rely on variations in the low pressure part of the engine. One example is to use a variable low pressure turbine operated in a heat exchanged cycle. However, this requires bulky and heavy heat exchangers with potential durability/reliability drawbacks.
Another example is to use variable turbine cooling, where the turbine cooling and associated losses are reduced during the cruise phase where maximum efficiency is sought. However, this concept can not give the full energy saving potential as the primary inefficiency of part power operation from lowering the temperature and pressure are not addressed.
Although the theory behind variable core flow gas turbine engines is very promising, there are still problems in practical implementation of this concept.
To achieve a variable core flow, U.S. Pat. No. 5,806,303 proposes the use of variable turbine stator blades for varying the high-pressure turbine area. Variable blades arranged close downstream of the combustion area are, however, exposed to a rough environment with very high temperatures. A drawback of this solution is therefore a questionable durability/reliability of the components involved.
Thus, there is still a need for improvements in this field.
It is desirable to provide a gas turbine engine, in particular a gas turbine engine for propulsion of an aircraft, that has a lower fuel consumption compared to conventional gas turbine engines.
The invention concerns a gas turbine engine comprising in serial flow relationship: a first compressor provided with at least one row of compressor blades distributed circumferentially around the first compressor; a combustion chamber; and a first turbine provided with at least one row of turbine blades distributed circumferentially around the first turbine, wherein the first compressor and the first turbine are rotationally rigidly connected by a first shaft.
The inventive gas turbine engine is characterized in that the first turbine is adapted to influence a gas flow rate through the gas turbine engine depending on a rotational speed of the first turbine, wherein the gas turbine engine further comprises means for controlling the rotational speed of the first turbine.
Thus, the inventive gas turbine engine makes it possible to vary the gas flow rate through the engine by controlling the rotational speed of the first turbine. An effect of decreasing the flow rate through the gas turbine engine is that the overall pressure ratio, i.e. the ratio between the pressure in the combustion chamber and the pressure at the main air inlet of the gas turbine engine, will increase. The invention thus decouples the pressure ratio and the flow rate from each other so that the gas turbine engine can be used to make a variable cycle engine. The gas turbine engine, i.e. the engine core of e.g. a turbo-fan engine, can now be run e.g. at a high flow rate and medium-high pressure ratio, in order to provide maximum power, or at a low flow rate and high pressure ratio, in order to provide maximum efficiency (i.e. minimum fuel consumption) at part power.
In conventional gas turbine engines both the gas flow rate and the pressure ratio increase with increasing amounts of fuel supplied to the combustion chamber. In the inventive gas turbine engine, it is possible to reach a higher pressure also at a lower level of fuel supply, i.e. it is possible to let the compressor operate at a higher operating line at part power.
That “the first turbine is adapted to influence a gas flow rate through the gas turbine engine depending on a rotational speed of the first turbine” means in particular that there is no inlet stator associated with the first turbine.
Normally, such a stator is considered to form part of the turbine and is positioned upstream of the turbine rotor for swirling the gas flow into the turbine rotor so as to increase the power output of the turbine. Turbine inlet stators are commonly arranged so that the flow is choked over a large part of the engine operating range. This simplifies the flow design of engine components in the compressor and the combustor. The flow rate of a choked gas flow can not be influenced by downstream variations, such as a varying rotation speed of a downstream turbine rotor. In contrast, the turbine of the invention allows the gas to flow in a more or less axial direction into the turbine rotor. The flow rate of an axially directed gas flow can be influenced to a higher degree by the rotational speed of the turbine than a swirling (non-choked) flow.
In an advantageous embodiment of the invention the means for controlling the rotational speed of the first turbine comprises a variable gas flow guiding means arranged upstream of at least one row of compressor blades of the first compressor, wherein the variable gas flow guiding means is adapted to guide the gas flow such as to influence the rotational speed of the first compressor. This way also the rotational speed of the first turbine is influenced since they are rotationally rigidly connected.
For instance, the guiding means can be used to go from a slight co-swirl to a slight counter-swirl of the inlet flow to the at least one row of compressor blades of the first compressor. This means that the rotational speed of the first compressor, and thus of the first shaft and the first turbine, will decrease. Due to the arrangement of the first turbine, e.g. the absence of a conventional turbine inlet stator, also the gas flow rate through the engine will decrease.
Preferably, the variable gas flow guiding means comprises a set of variable gas flow guide vanes. Guide vanes of this type are known as such and are therefore easy to adapt to the invention.
That “the variable gas flow guiding means is adapted to guide the gas flow such as to influence the rotational speed of the first compressor” means that said means is capable of influencing the degree of swirl of the gas that, during operation of the engine, flows through the first compressor, or at least of the gas that flows towards the at least one row of compressor blades positioned downstream of the guiding means. A skilled person in the art realizes that this, for instance, means that there should not be any other components arranged between the guiding means and the compressor blades that would significantly reduce the guide means' capability of influencing the swirl of the gas that flows towards these compressor blades. This capability is in some ways similar to the capability of conventional variable compressor guide vanes. Such conventional guide vanes are, however, mainly used for controlling the compressor stability, whereas the guiding means of the present invention is controlled in a different way for controlling the rotational speed of the first compressor.
In an advantageous embodiment of the invention the at least one row of turbine blades of the first turbine is arranged longitudinally adjacent to the combustion chamber. This means that the turbine rotor is arranged directly downstream of the combustion chamber and further clarifies that no stator vanes (nozzle blades) or similar gas deflecting components are arranged in a region downstream of the combustion chamber and upstream of the first turbine. Thus, the gas turbine engine is arranged to avoid any significant deflection of a gas flow that, during operation of the gas turbine engine, flows in this region.
In an advantageous embodiment of the invention the gas turbine engine comprises a second compressor and a second turbine rotationally rigidly connected by a second shaft, wherein the second compressor is arranged upstream of the combustion chamber, wherein the second shaft is arranged concentrically in relation to the first shaft and wherein the second turbine is positioned downstream of the first turbine. This way the efficiency and power output of the gas turbine engine can be increased. Preferably, the second shaft is arranged to rotate in an opposite direction in relation to the first shaft to increase the efficiency.
In an advantageous embodiment of the invention the second compressor is arranged upstream of the first compressor. Such a design makes it possible to reduce component stress and to use a simpler bearing arrangement.
In an advantageous embodiment of the invention the first compressor comprises a case arranged to rotate around the second compressor, which case is provided with a plurality of rows of compressor blades protruding in an inward direction towards the second compressor, wherein the first compressor and the second compressor overlap. Such a counter-rotating compressor makes it possible to provide a higher pressure ratio with fewer airfoils.
In an advantageous embodiment of the invention the second compressor extends in an axial direction over a longer distance than the first compressor such that the first and second compressors overlap only partly. Preferably, the second compressor has a plurality of rows of compressor blades positioned upstream of the most upstream row of compressor blades of the first compressor. A fully overlapping counter-rotating compressor is difficult to run at part speed since only one row of variable guide vanes can be provided, i.e. at the inlet. By arranging the counter-rotating compressor in a partly overlapping manner it becomes possible to provide the compressor with further rows of (non-inlet) variable guide vanes which is important for the stability of the compressor if a higher pressure ratio is desired.
In such a counter-rotating compressor, the variable gas flow guiding means is preferably positioned upstream of the first compressor and downstream of said plurality of rows of compressor blades positioned upstream of the most upstream row of compressor blades of the first compressor.
In an advantageous embodiment of the invention a first rotor arrangement is attached to a first carrying frame positioned downstream of the first compressor and upstream of the combustion chamber, wherein the first rotor arrangement comprises the first compressor, the first shaft and the first turbine. Where overlapping compressor are used the first rotor arrangement is preferably also attached to a second carrying frame that is positioned upstream of the first compressor. This allows for a working bearing arrangement without intershaft bearings. The variable gas flow guiding means is preferably attached to the second carrying frame.
Preferably, the radius of a gas turbine engine core is less along a distance where the first compressor and the second compressor overlap than upstream of the first compressor. Preferably, the radius of the gas turbine engine is reduced, as seen in a downstream direction, at or around the second carrying frame. Such a design allows the stator-rotor stages upstream of the first compressor to run sufficiently fast, i.e. to run at a sufficiently high Mach number, to provide a high pressure ratio per stage, while the downstream stages, where counter-rotation is applied, are allowed to run sufficiently slow to avoid excessive shock losses.
In an advantageous embodiment of the invention the turbine blades of the first turbine have an outlet blade angle of at least 60°. The outlet blade angle should be sufficiently large for providing a sufficiently small area between the blades to reduce the flow rate sufficiently when the turbine is retarded. Preferably, the turbine blades of the first turbine have a camber of around 45° or lower. To maximize the gas flow when the first turbine is only lightly loaded, e.g. by an appropriate setting of the variable inlet guide vanes of the compressor, the profile of the turbine blades should have a low amount of camber.
In an advantageous embodiment of the invention the means for controlling the rotational speed of the first turbine comprises an arrangement for taking out mechanical power from the first shaft. This can be used as an alternative or in combination with the variable gas flow guiding means. Preferably, the arrangement for taking out mechanical power comprises an electric generator connected to the first shaft.
The invention also concerns a method for controlling a gas flow rate in an axial flow gas turbine engine. The inventive method comprises the step of: controlling a rotational speed of a turbine, wherein the turbine is adapted to influence the gas flow rate through the gas turbine engine depending on the rotational speed of the turbine.
In an advantageous embodiment of the inventive method it comprises the step of: adjusting a variable gas flow guiding means arranged upstream of at least one row of compressor blades of a compressor that is rotationally rigidly connected to the turbine. Preferably, it comprises the step of: decreasing the rotational speed of the turbine by adjusting the variable gas flow guiding means such as to increase a counter-swirl, or decrease a co-swirl, of the gas flow.
In an advantageous embodiment of the inventive method it comprises the step of: taking out mechanical power from a shaft that is rotationally rigidly connected to the turbine.
The invention also concerns an axial flow gas turbine engine comprising a first rotor arrangement comprising a first compressor, and a second rotor arrangement comprising a second compressor, wherein the first and second rotor arrangements are positioned concentrically and are arranged to rotate in opposite directions, and wherein the first compressor comprises a case arranged to rotate around the second compressor, which case is provided with a plurality of rows of compressor blades protruding in an inward direction towards the second compressor, wherein the first compressor and the second compressor overlap. In the invention the second compressor extends in an axial direction of the gas turbine engine over a longer distance than the first compressor such that the first and second compressors overlap only partly. As mentioned above, such a partly overlapping, counter-rotating compressor arrangement makes it possible to provide the compressor with further rows of (non-inlet) variable guide vanes which is important for the stability of the compressor if a higher pressure ratio is desired. A gas turbine engine provided with such a compressor arrangement is useful in various applications.